A turbine engine used to power aircraft typically includes a compressor section for compressing air, a combustor section for mixing the compressed air with fuel and burning it to generate hot gases, a turbine section in which energy is extracted from the hot gases of combustion to operate the compressor and otherwise provide the aircraft and an exhaust section in which the remaining hot gases of combustion of reduced energy pass to provide thrust which propels the aircraft. A turbofan turbine engine further includes a large fan typically located in front of the compressor. In the case of a front fan, the fan is driven by a second turbine, typically the low pressure turbine, located in the turbine section of the engine behind the high pressure turbine. In a low bypass turbofan turbine engine, typically used for supersonic flight, most of the air passing the fan enters into the compressor and an augmenter is added to supplement thrust. In a high bypass turbofan turbine engine, typically used for subsonic flight, more air flows around (bypasses) the engine, contributing to thrust while reducing specific fuel consumption. The present invention is directed to the combustor section of the turbine engine. Although the present invention finds immediate use in a high bypass turbofan turbine engine, specifically the GE-90, currently manufactured by General Electric Company, it also has applicability to other turbine engines including other high bypass turbine engines and low bypass turbine engines,
As noted, in the combustor, fuel is mixed with compressed air and burned. The combustor is located in the engine aft of the compressor. The compressed air leaving the compressor passes into the combustor. The compressed air exits the compressor at a high velocity, with low static pressure. The diffuser at the aft end of the compressor reduces the air velocity, increasing the static pressure of the air. High static pressure and low velocity at the combustor entrance improves mixing in the combustion chamber. Fuel is metered into a swirl cup before it passes into the combustor.
The combustor is an assembly which typically consists of several individual pieces. The dome assembly establishes a bluff body to stabilize the flame. The inner and outer cowls and the inner and outer liners are typically bolted to the dome at its inside and outside diameter. The bolts and locking nuts typically are tack welded to prevent them from separating. The dome structure consists of a single spectacle plate, a die formed sheet metal part. A plurality of swirl cups are included in the dome assembly for swirling the fuel as it is mixed with the air. An igniter is provided to ignite the fuel-air mixture and burn it. The inner and outer combustor liners are typically fabricated from individual machined forgings or castings and are welded or bolted together.
The combustor liners form a substantially closed duct, typically annular, that channels the hot gases of combustion to the high pressure turbines. These combustor liners must withstand the temperature of combustion which can be 3000° F. or higher. To survive these conditions, exotic high temperature alloys are utilized. Each combustor is located within an inner and an outer structure. Cooling air flows along the exterior surfaces of the combustor in the annular region between the inner liners and the inner structure and outer liners and outer structure. To further assist in the durability of the combustor structure, cooling air is introduced through a plurality of cooling holes in the combustor liner. In addition, dilution air is introduced through several rows of holes in the liners to further mix the reacting gases. Additionally, the inner surface of the liners, the surface adjacent to the hot combustion gases, is coated with a thermal barrier coating systems. Many thermal barrier coating systems exist, and any of these can be used. The thermal barrier coating systems typically include a metallic bond coat and a ceramic top coat to insulate the surface of the liners from the hot gases of combustion.
Combustor liners are formed from a single panel or a plurality of combustor liner panels that are welded together. The liner is attached to the dome assembly by a forward bolt flange. Recently, inner liners after usage in an engine have experienced cracking in two regions of the aft panel assembly, the multi-hole cooling area of the liner. The cracks in the multi-hole cooling area are in line with the fuel cups located in the dome and are typically associated with combustion hot streaks. The cracks in the aft panel flange region are in the area of the aft lip of the liner, located in the portion of the liner adjacent the high pressure nozzle are in line with the high pressure turbine stage 1 nozzles. The cracks may be due to low cycle fatigue caused by large circumferential temperature gradients that reduce the liner low cycle fatigue life in that region. In the past, these cracks have been repaired by welding or other alloy repair process. Weld repair requires further processing to restore dimensional and operational requirements that become economically excessive at some level of cracking. Additionally, the number and location of cracks makes it such that the weld repair is not feasible because hole restoration in the multi-hole cooling area is extensive. At the aft panel flange region, crack repair can result in aft lip distortion and shrinkage. This is unacceptable as the exit throat diameter of the liner is a critical dimension. In the past, the other alternative was to discard the cracked liner and replace it with a new liner. Typically, the liner is made from GTD-222, which has a nominal composition, in weight percent, of about 22.5 chromium, about 14.0 cobalt, about 2.3 titanium, about 1.2 aluminum, about 2.0 tungsten, about 0.8 columbium, and about 1.0 tantalum, with the balance being nickel and incidental impurities, but the liner is not so limited, and other oxidation resistant high temperature alloys may also be used. GTD-222 is an expensive alloy, as are other high temperature, oxidation resistant alloys that may be used in this application, and the processing involved in its manufacture adds to the value of the liner. Thus, it is desirable to develop a method for repairing these liners, so as to salvage as much of the expensive liner as possible and limiting replacement material as much as possible.
A method for repairing the liners is set forth in U.S. Pat. No. 6,568,079 to Farmer et al. and assigned to the assignee of the present application, issued May 27, 2003. This patent sets forth an acceptable method for replacing damaged panels in a combustor. However, the repair method does not address repair of a portion of the aft panel assembly, which is discarded and replaced. The aft panel assembly includes the aft panel, the aft lip, the seal flange area and aft panel support leg. The support leg extends radially inward in the engine across the annular region between the inner liner and the structure and is bolted to the inner structure. Although portions of the aft panel assembly experience crack damage, the aft panel support leg typically is not affected. What is needed is a method for repairing the aft panel of an inner liner by replacing only the damaged portion of the aft panel assembly, without necessitating the scrapping of the aft panel support leg.